This invention relates to generally to orbiting satellites and, more particularly, to packaging and cooling techniques for electronic modules carried on satellites. Satellites consist of various modules, which, largely for reasons of convenience and economy, are typically designed and manufactured by separate entities within a company, or by separate companies entirely. For example, a commonly accepted distinction is between an electronics module, such as a payload module, designed to perform a specific function in space, and the structure of a spacecraft designed to perform a support function for one or more modules that it carries. Conventionally, payload modules are constructed to be housed in six-sided metal payload boxes, which are secured to the spacecraft, usually by bolting down onto part of the spacecraft structure.
Each payload box typically houses heat-producing electronics components, which must be maintained below a maximum operating temperature to ensure that they operate for a desired operating life without defects. In conventional satellite design, each box provides a thermal path from the heat-producing components to a metal baseplate on one face of the box. When the payload box is integrated into the satellite, this baseplate is secured to a heat-conducting structure on the spacecraft, including another metal plate referred to as a doubler plate, which further spreads the thermal path laterally, a honeycomb structure with embedded heat pipes, and a heat radiator panel. Heat from the payload components conducts through the baseplate and is spread by the doubler plate, the honeycomb structure and the heat pipes across a larger area of the radiator panel, from which the heat is radiated into space.
There are two principal drawbacks to the conventional satellite structure described above. First, the thermal path between the heat-producing components and the radiator panel includes a number of thermal resistance components that together result in a lower radiated power from the panel and a higher operating temperature of the components. In addition, the spacecraft structure that contributes significantly to the thermal path resistance, also contributes to the total mass and the cost of the spacecraft. However, so long as the baseplate interface between payload boxes and the spacecraft structure is mandated by convention, these thermal resistance components cannot be eliminated or easily reduced. The second drawback of the conventional satellite structure as described is that any modification of components in a payload box can be accomplished only by completely removing the box from the spacecraft. Replacement of components or circuit cards during integration and testing of the satellite is, therefore, a time-consuming and expensive procedure. Similarly, replacement of faulty components immediately before launch of the satellite, or while the satellite is in orbit, is equally difficult and time-consuming.
Ideally, a satellite should be designed to minimize the thermal resistance between heat-producing components and a heat-radiating panel, to minimize spacecraft mass, and to provide convenient access to payload components for maintenance and replacement. As will become apparent from the description that follows, the present invention achieves goals.
The present invention resides in a satellite structure in which each payload box has five conventional faces and a sixth face that serves as part of a direct radiator panel structure and as an access panel for maintenance of components housed in the box.
Briefly, and in general terms, the invention may be defined as a spacecraft module comprising one or more heat-generating electronics components; a metal box enclosing the electronics components and having one face that forms an opening to provide access to the box; and a heat radiating panel, including a heat-conductive backing plate for installation over the opening in the box, with the backing plate in a direct heat-conductive relationship with the electronics components. The heat radiating panel is removable to facilitate access to the electronics components, and the thermal resistance between the electronics components and the heat radiating panel is minimized by the direct transmission of heat to the radiator, without any intervening heat-conducting structures. Therefore, both the mass and the cost of each module, and of the entire spacecraft, are significantly reduced. More specifically, the metal box has five solid faces and a sixth face that forms the opening to provide access to the box.
The structure of the invention may also be defined as a spacecraft module, comprising a spacecraft structure, including a cavity that presents an opening at an external surface of the structure; an electronics module box having four contiguous faces forming sidewalls, a fifth face adjoining the sidewalls and sixth face that includes an opening for access to the module box, wherein the module box is installed in the cavity of the spacecraft structure, with the opening to the module box approximately coplanar with the opening to the cavity in the spacecraft structure. The structure of the invention further comprises a plurality of electronics components mounted on circuit boards and installed in the module box; a supporting structure installed in the module box, to provide mechanical support for the circuit boards and a thermal path for heat generated in the electronics components; a removable radiator panel structure, including a thermally conductive backing plate; and a plurality of fasteners to attach the radiator panel structure to the spacecraft structure. When attached to the spacecraft, the radiator panel structure covers the opening to the cavity in the spacecraft structure and covers the module box opening, such that the backing plate of the radiator panel is placed in good thermal contact with the supporting structure in the module box. Heat generated by the electronics components is transmitted directly from the supporting structure to the radiator panel structure, and the thermal resistance between the supporting structure and the radiator panel structure is minimized. An important benefit is that access to the electronics components is easily effected by removal of the radiator panel structure.
The invention may also be defined as a method for reducing the mass of, and improving maintenance access to, a spacecraft module. The method comprises the steps of providing an electronics module box, with one face missing to define an opening for maintenance of electronics components enclosed therein; installing electronics components in the electronics module box; installing the electronics module box in a spacecraft, with the opening facing out; and installing a radiator panel over the opening in the box, wherein the panel includes a backing plate installed in thermal contact with the electronics components enclosed in the box. Spacecraft mass is significantly reduced by ensuring more direct contact between the electronics components and the radiator panel, and maintenance access is more easily effected by removal of the radiator panel. Accordingly, the method may further comprise the steps of removing the radiator panel; repairing electronics components in the box without removing the box from the spacecraft; and replacing the radiator panel.
It will be appreciated from the foregoing summary that the present invention represents a significant advance in the field of spacecraft architecture. In particular, the invention achieves significant reductions in mass and cost by eliminating components that are conventionally employed to couple an electronics module box to a spacecraft structure. Direct coupling of heat generated by electronics components to the radiator panel also facilitates access to the electronics components, by removal of the panel. Other aspects and advantages of the invention will become apparent from the following more detailed description, taken in conjunction with the accompanying drawings.